airfoil: drag and lift coefficients
I want to calculate the drag coefficient and lift coefficient on an airfoil (2d wing).When I go to Report>>Forces it shows the pressure coeficcient, total force and pressure force, which is something that I don't need. Also if i request to calculate the area of the wing it gives 0.20426972m2.On the reference values the area is 1.If I change the are here doesn't do anything...
What should I do?
Re: airfoil: drag and lift coefficients
Hi Pedro, I'm working on such a field like yours..I'm studing the Cl and Cd curve of 2D profiles dipending on different angle of attack (from 0 to 90)...I suggest you not to take care about the references value but just to go in Solver>>forces>>and set Cl and Cd, select the option PRINT and after the simulation is converged ( it depends on the criteria you require)you can see the value of Cd and Cl..
I suggest you also to follow the fluent tutorial file about airfoil.
If you want we can keep in contact to share ideas and result
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